EP2071152B1 - Diffuser of a turbomachine - Google Patents

Diffuser of a turbomachine Download PDF

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Publication number
EP2071152B1
EP2071152B1 EP08162167.4A EP08162167A EP2071152B1 EP 2071152 B1 EP2071152 B1 EP 2071152B1 EP 08162167 A EP08162167 A EP 08162167A EP 2071152 B1 EP2071152 B1 EP 2071152B1
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EP
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Prior art keywords
flow
turbomachine
diffuser
axis
straightening vanes
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EP08162167.4A
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German (de)
French (fr)
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EP2071152A2 (en
EP2071152A3 (en
Inventor
Patrice Commaret
Michel Desaulty
Romain Lunel
Pascale Rollet
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Safran Aircraft Engines SAS
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Safran Aircraft Engines SAS
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/08—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising at least one radial stage
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/42—Casings; Connections of working fluid for radial or helico-centrifugal pumps
    • F04D29/44—Fluid-guiding means, e.g. diffusers
    • F04D29/441—Fluid-guiding means, e.g. diffusers especially adapted for elastic fluid pumps
    • F04D29/444—Bladed diffusers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved

Definitions

  • the invention relates to a turbomachine. It is intended for any type of turbomachine, terrestrial or aeronautical, and more particularly to aircraft turbojets.
  • upstream and downstream are defined with respect to the normal flow direction of the gas (from upstream to downstream) through the turbomachine.
  • the axis of the turbomachine is called the axis of rotation of the rotor of the turbomachine.
  • the axial direction corresponds to the direction of the axis of the turbomachine, and a radial direction is a direction perpendicular to this axis.
  • an axial plane is a plane containing the axis of the turbomachine and a radial plane is a plane perpendicular to this axis.
  • the adjectives inner / inner and outer / outer are used with reference to a radial direction so that the inner / inner (ie radially inner) part of an element is closer to the axis of the turbomachine than the outer / outer (ie radially outer) part of the same element.
  • a compressor having a downstream centrifugal stage usually comprises a rotary impeller.
  • This impeller comprises a series of blades driven in rotation and is designed to accelerate the gas that passes through it.
  • the diffuser has an annular space surrounding the impeller.
  • the diffuser serves to reduce the speed of the gas leaving the impeller and thereby increase its static pressure.
  • the diffusers may be paddle type or duct type.
  • These two types of diffuser comprise, in general, a radially oriented annular upstream portion which has a series of diffusion passages connected to the output of the compressor to recover the accelerated gas leaving it.
  • These diffusion passages have a section which increases gradually from upstream to downstream in order to diffuse the flow of gas leaving the compressor.
  • the paddle-type diffusers use a series of circularly spaced vanes forming between them these diffusion passages.
  • these passages are constituted by pipe elements or ducts, formed for example between two opposed opposite plates.
  • the diffusers Downstream of this upstream part, the diffusers comprise, in general, an angled annular intermediate portion for bending the flow path of the diffuser and reducing the flow of gas towards the combustion chamber.
  • the diffusers comprise, in general, an annular downstream part comprising a series of circularly spaced rectifying blades, making it possible to straighten the gas flow and thus reduce or eliminate the circumferential swirling of the flow of gas exiting the diffusion passages, before this flow enters the combustion chamber.
  • the centers of the injection orifices of the combustion chamber are distributed around the axis of the turbomachine, on a circle of radius R1, while the average radius R2 of the downstream part of the diffuser is greater than the radius R1.
  • the downstream part of the diffuser follows the line of the outer casing of the chamber and is directed towards the external bypass zone of the chamber (ie the zone of passage between the chamber and the outer casing).
  • the average axis of the current stream at the outlet of the downstream portion of the diffuser is parallel to the mean axis of the bypass flow external of the room.
  • the injection systems and the internal bypass zone are then fed by a secondary flow deviated from the main flow, this deviation generating a significant loss of pressure (ie a loss of pressure) between the outlet of the diffuser and the upstream of the injection system and between the outlet of the diffuser and the internal bypass zone.
  • downstream portion of the diffuser is inclined relative to the axis of the turbomachine, in the direction of the combustion chamber, so that, in a section plane containing the axis of the turbomachine, the mean axis of the vein current output of the downstream portion of the diffuser, passes through the chamber bottom, between the maximum radius and the minimum radius of the chamber bottom.
  • the stream of current is defined as the envelope which delimits the flow space of the gas, and therefore the flow of gas.
  • this vein is delimited by the internal contour of this downstream part.
  • Such inclination of the downstream part of the diffuser relative to the axis of the turbomachine towards the chamber bottom is an improvement because it reduces the pressure drop between the outlet of the diffuser and the upstream of the system. injection, by feeding these systems more directly. It also makes it possible to supply more symmetrically gas to the external and internal bypass zones of the chamber, and also to better supply gas to the internal bypass zone. In addition, the gas supply of the different intake channels of each injection system is also more homogeneous.
  • the invention aims to improve the latter type of turbomachine, to further improve the supply of injection systems.
  • the subject of the invention is a turbomachine, of the last aforementioned type (ie with a downstream part of the diffuser inclined with respect to the axis of the turbomachine, in the direction of the chamber bottom), according to the object of the independent claim 1.
  • the inventors have found that the straightening vanes of the downstream part of the diffuser induce disturbances in the gas flow downstream of their trailing edge (we speak of wakes), and these disturbances have a detrimental effect on the gas supply of the injection systems.
  • this gas supply is less symmetrical around the injection axis of these systems.
  • the invention therefore proposes to distance sufficiently the injection orifices (and therefore the injection systems) from the trailing edge of the straightening vanes, so that these orifices are in an area of slight disturbance or even zero disturbance, and thus improve the supply of the injection systems.
  • the invention there is a better supply of the injection system (in particular more symmetrical about the injection axis), which allows, in particular, to improve the stability of the low-speed combustion, better control of the temperature profile at the outlet of the chamber and to limit the risk of non-stationary coupling between the outlet flow of the diffuser and the combustion.
  • the injection orifices must not be too far from the straightening vanes to limit the burst pressure drop between the outlet of the diffuser and the upstream of the injection system.
  • the distance, in abscissa curvilinear, along a current line between the middle of the current stream at the trailing edge of the straightening vanes and said center is less than or equal to 9 times the height of said stream of current at the trailing edge of the straightening vanes.
  • the number of straightening vanes is greater than the number of injection systems.
  • the number of straightening vanes is at least 4 times greater than the number of injection systems. The greater the number of straightening vanes, the more turbulent structures induced by these vanes are numerous (but of smaller size), and the more the invention is of interest since it makes it possible to dissipate these turbulent structures on the recommended distance between the trailing edge of the straightening vanes and the injection ports.
  • the Figures 1 to 3 represent an example of a turbojet engine in axial half-section along a section plane containing the rotation axis A of the turbojet rotor.
  • the gas passing through the turbojet is air.
  • the turbojet engine comprises a high pressure compressor 10 whose part swallows (visible on the Figures 1 to 3 ) is composed of a centrifugal stage, an annular diffuser connected downstream of the compressor 10, this diffuser opening into a space 30 surrounding an annular combustion chamber 40.
  • This space 30 is delimited by an outer casing 32 and a concentric inner casing 34.
  • the combustion chamber 40 is supported by fixing clamps connected to the housings 32 and 34.
  • the centrifugal compressor comprises a rotary impeller 12.
  • This impeller 12 comprises a series of moving blades 14, driven in rotation.
  • the impeller 12 is designed to accelerate the air passing through it and thereby increase the kinetic energy of this air.
  • the diffuser 20 has an annular space surrounding the impeller.
  • the diffuser 20 serves to reduce the speed of the air leaving the impeller and thereby increase its static pressure.
  • the diffuser 20 of the figures is of the paddle type.
  • This diffuser 20 has a radially oriented annular upstream portion 21 which has a series of diffusion passages 22 connected to the output of the compressor 10 to recover the accelerated air coming out of the impeller 12.
  • These diffusion passages 22 have a section which progressively increases from upstream to downstream in order to diffuse the air flow leaving the impeller.
  • These diffusion passages 22 are formed by a series of blades 23 spaced apart in a circular manner. At the inlet of the upstream portion 21, these blades 23 are close to each other. These vanes 23 circumferentially move away from each other as they approach the outlet of the upstream portion 21.
  • the diffuser 20 Downstream of the upstream portion 21, the diffuser 20 comprises an annular intermediate portion 24 bent, to curve the flow path of the diffuser and reduce the flow of air towards the combustion chamber 40.
  • the diffuser 20 Downstream of this intermediate portion 24, the diffuser 20 comprises an annular downstream portion 25 comprising a series of circumferentially spaced rectifying vanes 26 for eliminating or limiting the circumferential swirling of the outgoing air flow from the diffusion passages 22, before this airflow enters the space 30.
  • the figure 5 represents the straightening vanes 26 in section, in a sectional plane perpendicular to the mean axis X of the stream of current at the outlet of the downstream part 25 of the diffuser. As seen in this figure, the straightening vanes 26 extend radially between the inner and outer walls of the downstream part 25 of the diffuser.
  • the combustion chamber 40 comprises an annular inner wall 42, an annular outer wall 43 and an annular chamber bottom 41 disposed between said walls 42, 43, in the upstream region of said chamber.
  • This chamber bottom 41 has injection orifices 44 distributed circumferentially about the axis A.
  • Injection systems 45 are mounted on the chamber bottom, through said injection ports 44 (due to a system injection 45 by injection port 44). These injection systems 45 make it possible to inject the air / fuel mixture which is burned into the combustion chamber 40.
  • the fuel of this mixture is fed to the injection systems 45 via a fuel feed pipe 46 running through the space 30.
  • the combustion chamber 40, its housings and its close environment are commonly referred to as the combustion chamber module.
  • the combustion chamber 40 is inclined relative to the axis A of the turbojet engine of an acute angle B (non-zero).
  • This acute angle B is, the smaller the axial size of the combustion chamber module.
  • the downstream portion 25 of the diffuser 20 is inclined relative to the axis A of the turbojet engine, in the direction of the combustion chamber, so that, in a section plane containing the axis of the turbojet, the average axis X of the stream of current at the outlet of the downstream part 25 of the diffuser, passes through the chamber bottom 41, between the maximum radius R and the minimum radius r of the chamber bottom 41.
  • the chamber bottom 41 being annular and centered on the axis A, the r and r rays extend from the axis A in a radial direction.
  • Figures 1 to 3 represent three examples of combustion chamber 40 for which the aforementioned inclination criterion is respected.
  • the mean axis X of the stream of current at the outlet of the downstream part 25 of the diffuser passes through the chamber bottom 41 close to the maximum radius R.
  • the mean axis X of the current stream at the outlet of the downstream part 25 of the diffuser passes through the chamber bottom 41 near the minimum radius r.
  • the mean axis X of the stream of current at the outlet of the downstream part 25 of the diffuser passes through the chamber bottom 41 at the center C of an injection orifice 44.

Description

L'invention concerne une turbomachine. Elle se destine à tout type de turbomachine, terrestre ou aéronautique, et plus particulièrement aux turboréacteurs d'avion.The invention relates to a turbomachine. It is intended for any type of turbomachine, terrestrial or aeronautical, and more particularly to aircraft turbojets.

Dans la présente demande, l'amont et l'aval sont définis par rapport au sens d'écoulement normal du gaz (de l'amont vers l'aval) à travers la turbomachine. Par ailleurs, on appelle axe de la turbomachine, l'axe de rotation du rotor de la turbomachine. La direction axiale correspond à la direction de l'axe de la turbomachine, et une direction radiale est une direction perpendiculaire à cet axe. De même, un plan axial est un plan contenant l'axe de la turbomachine et un plan radial est un plan perpendiculaire à cet axe. Enfin, sauf précision contraire, les adjectifs intérieur/interne et extérieur/externe sont utilisés en référence à une direction radiale de sorte que la partie intérieure/interne (i.e. radialement intérieure) d'un élément est plus proche de l'axe de la turbomachine que la partie extérieure/externe (i.e. radialement extérieure) du même élément.In the present application, upstream and downstream are defined with respect to the normal flow direction of the gas (from upstream to downstream) through the turbomachine. Furthermore, the axis of the turbomachine is called the axis of rotation of the rotor of the turbomachine. The axial direction corresponds to the direction of the axis of the turbomachine, and a radial direction is a direction perpendicular to this axis. Similarly, an axial plane is a plane containing the axis of the turbomachine and a radial plane is a plane perpendicular to this axis. Finally, unless otherwise stated, the adjectives inner / inner and outer / outer are used with reference to a radial direction so that the inner / inner (ie radially inner) part of an element is closer to the axis of the turbomachine than the outer / outer (ie radially outer) part of the same element.

L'invention concerne une turbomachine du type comprenant :

  • une chambre de combustion annulaire avec un fond de chambre présentant des orifices d'injection, dans lesquels sont montés des systèmes d'injection ;
  • un compresseur comportant un étage aval centrifuge ; et
  • un diffuseur annulaire permettant de diffuser le courant de gaz sortant du compresseur et de diriger ce courant de gaz vers lesdits systèmes d'injection.
The invention relates to a turbomachine of the type comprising:
  • an annular combustion chamber with a chamber bottom having injection ports, in which injection systems are mounted;
  • a compressor having a centrifugal downstream stage; and
  • an annular diffuser for diffusing the stream of gas leaving the compressor and directing the gas stream to said injection systems.

Un compresseur comportant un étage aval centrifuge comprend habituellement un impulseur rotatif. Cet impulseur comprend une série d'aubes entraînées en rotation et est réalisé de façon à accélérer le gaz qui le traverse.A compressor having a downstream centrifugal stage usually comprises a rotary impeller. This impeller comprises a series of blades driven in rotation and is designed to accelerate the gas that passes through it.

Le diffuseur présente un espace annulaire entourant l'impulseur. Le diffuseur sert à réduire la vitesse du gaz quittant l'impulseur et, de ce fait, à accroître sa pression statique. Les diffuseurs peuvent être du type à aubes ou du type à conduits.The diffuser has an annular space surrounding the impeller. The diffuser serves to reduce the speed of the gas leaving the impeller and thereby increase its static pressure. The diffusers may be paddle type or duct type.

Ces deux types de diffuseur comprennent, en général, une partie amont annulaire orientée radialement qui présente une série de passages de diffusion raccordés à la sortie du compresseur pour récupérer le gaz accéléré sortant de celui-ci. Ces passages de diffusion ont une section qui s'accroît progressivement de l'amont vers l'aval afin de diffuser le courant de gaz sortant du compresseur. Les diffuseurs du type à aubes utilisent une série d'aubes espacées circulairement et formant entre elles ces passages de diffusion. Dans les diffuseurs du type à conduits, ces passages sont constitués par des éléments de tuyaux ou conduits, formés par exemple entre deux plaques opposées jointes.These two types of diffuser comprise, in general, a radially oriented annular upstream portion which has a series of diffusion passages connected to the output of the compressor to recover the accelerated gas leaving it. These diffusion passages have a section which increases gradually from upstream to downstream in order to diffuse the flow of gas leaving the compressor. The paddle-type diffusers use a series of circularly spaced vanes forming between them these diffusion passages. In the duct type diffusers, these passages are constituted by pipe elements or ducts, formed for example between two opposed opposite plates.

En aval de cette partie amont, les diffuseurs comprennent, en général, une partie intermédiaire annulaire coudée, pour courber le trajet d'écoulement du diffuseur et ramener l'écoulement de gaz en direction de la chambre de combustion.Downstream of this upstream part, the diffusers comprise, in general, an angled annular intermediate portion for bending the flow path of the diffuser and reducing the flow of gas towards the combustion chamber.

En aval de cette partie intermédiaire, les diffuseurs comprennent, en général, une partie aval annulaire comprenant une série d'aubes de redressement circulairement espacées, permettant de redresser l'écoulement de gaz et ainsi de réduire ou de supprimer le tourbillonnement circonférentiel de l'écoulement de gaz sortant des passages de diffusion, avant que cet écoulement n'entre dans la chambre de combustion.Downstream of this intermediate part, the diffusers comprise, in general, an annular downstream part comprising a series of circularly spaced rectifying blades, making it possible to straighten the gas flow and thus reduce or eliminate the circumferential swirling of the flow of gas exiting the diffusion passages, before this flow enters the combustion chamber.

Généralement, les centres des orifices d'injection de la chambre de combustion sont repartis autour de l'axe de la turbomachine, sur un cercle de rayon R1, tandis que le rayon moyen R2 de la partie aval du diffuseur est supérieur au rayon R1.Generally, the centers of the injection orifices of the combustion chamber are distributed around the axis of the turbomachine, on a circle of radius R1, while the average radius R2 of the downstream part of the diffuser is greater than the radius R1.

Dans certaines turbomachines connues, la partie aval du diffuseur suit la ligne du carter extérieur de la chambre et est dirigée vers la zone de contournement externe de la chambre (i.e. la zone de passage entre la chambre et le carter extérieur). En d'autres termes, dans un plan de coupe contenant l'axe de la turbomachine, l'axe moyen de la veine de courant en sortie de la partie aval du diffuseur, est parallèle à l'axe moyen de l'écoulement de contournement externe de la chambre. Cette solution est insatisfaisante car l'ensemble de l'écoulement de gaz principal sortant du diffuseur contourne extérieurement la chambre de combustion avant de se répartir entre l'écoulement externe et l'écoulement alimentant le fond de chambre et la zone de contournement interne de la chambre (i.e. la zone de passage entre la chambre et le carter intérieur). Les systèmes d'injection et la zone de contournement interne sont alors alimentés par un écoulement secondaire dévié de l'écoulement principal, cette déviation engendrant une perte de charge (i.e. une perte de pression) notable entre la sortie du diffuseur et l'amont du système d'injection et entre la sortie du diffuseur et la zone de contournement interne.In certain known turbomachines, the downstream part of the diffuser follows the line of the outer casing of the chamber and is directed towards the external bypass zone of the chamber (ie the zone of passage between the chamber and the outer casing). In other words, in a cutting plane containing the axis of the turbomachine, the average axis of the current stream at the outlet of the downstream portion of the diffuser, is parallel to the mean axis of the bypass flow external of the room. This solution is unsatisfactory because the entire flow of the main gas leaving the diffuser bypasses the combustion chamber outside before distributing between the external flow and the flow supplying the chamber bottom and the internal bypass zone of the chamber (ie the zone of passage between the chamber and the inner casing). The injection systems and the internal bypass zone are then fed by a secondary flow deviated from the main flow, this deviation generating a significant loss of pressure (ie a loss of pressure) between the outlet of the diffuser and the upstream of the injection system and between the outlet of the diffuser and the internal bypass zone.

Les conséquences fonctionnelles d'une telle perte de charge sont les suivantes :

  • La perte de charge importante entre la sortie du diffuseur et le système d'injection doit être compensée lors de la conception de la turbomachine par une augmentation globale de la perte de charge du module, entre la sortie du diffuseur et la sortie de la chambre de manière à conserver une chute de pression suffisante à la traversée du système d'injection pour assurer le mélange air-carburant et la combustion. Cette augmentation de la perte de charge du module entraîne une augmentation de la consommation de carburant.
  • L'alimentation en gaz entre la zone de contournement externe et la zone de contournement interne de la chambre est fortement dissymétrique (les jets de gaz primaires et de dilution sont plus pénétrant en externe qu'en interne), ce qui rend plus difficile la maîtrise du profil de température en sortie de chambre.
  • Une mauvaise alimentation en gaz de la zone de contournement interne entraîne une diminution des vitesses d'écoulement des gaz dans les dispositifs de refroidissement de la paroi interne de la chambre de combustion, ce qui réduit les coefficients d'échange convectif et donc l'efficacité globale de ce refroidissement.
  • Une mauvaise alimentation en gaz de la zone de contournement interne entraîne un taux de surpression réduit diminuant l'efficacité du refroidissement du distributeur de turbine situé en aval de la chambre.
The functional consequences of such a loss of load are as follows:
  • The significant pressure drop between the outlet of the diffuser and the injection system must be compensated during the design of the turbomachine by an overall increase in the pressure drop of the module, between the outlet of the diffuser and the exit of the chamber. in order to maintain a sufficient pressure drop across the injection system to ensure air-fuel mixing and combustion. This increase in the pressure drop of the module leads to an increase in fuel consumption.
  • The gas supply between the outer bypass zone and the internal bypass zone of the chamber is strongly asymmetrical (the primary and dilution gas jets are more penetrating externally than internally), which makes it more difficult to control the temperature profile at the chamber outlet.
  • Poor gas supply to the internal bypass zone results in a decrease in gas flow velocities in the cooling devices of the inner wall of the combustion chamber, which reduces the convective exchange coefficients and thus the efficiency overall of this cooling.
  • Poor gas supply to the internal bypass zone results in a reduced overpressure rate decreasing the cooling efficiency of the turbine distributor located downstream of the chamber.

Pour éviter ces inconvénients, dans d'autres turbomachines connues, comme celle du document FR 2372965 , la partie aval du diffuseur est inclinée par rapport à l'axe de la turbomachine, en direction de la chambre de combustion, de sorte que, dans un plan de coupe contenant l'axe de la turbomachine, l'axe moyen de la veine de courant en sortie de la partie aval du diffuseur, passe par le fond de chambre, entre le rayon maximum et le rayon minimum du fond de chambre. La veine de courant se définit comme étant l'enveloppe qui délimite l'espace d'écoulement du gaz, et donc le courant de gaz. Au niveau de la partie aval du diffuseur, cette veine est délimitée par le contour intérieur de cette partie aval.To avoid these disadvantages, in other known turbomachines, such as that of the document FR 2372965 the downstream portion of the diffuser is inclined relative to the axis of the turbomachine, in the direction of the combustion chamber, so that, in a section plane containing the axis of the turbomachine, the mean axis of the vein current output of the downstream portion of the diffuser, passes through the chamber bottom, between the maximum radius and the minimum radius of the chamber bottom. The stream of current is defined as the envelope which delimits the flow space of the gas, and therefore the flow of gas. At the level of the downstream part of the diffuser, this vein is delimited by the internal contour of this downstream part.

Une telle inclinaison de la partie aval du diffuseur par rapport à l'axe de la turbomachine, en direction du fond de chambre, constitue un perfectionnement car elle permet de réduire la perte de charge entre la sortie du diffuseur et l'amont du système d'injection, en alimentant plus directement ces systèmes. Elle permet également d'alimenter en gaz, de façon plus symétrique, les zones de contournement externe et interne de la chambre, et aussi de mieux alimenter en gaz la zone de contournement interne. Par ailleurs, l'alimentation en gaz des différents canaux d'admission de chaque système d'injection est également plus homogène.Such inclination of the downstream part of the diffuser relative to the axis of the turbomachine towards the chamber bottom is an improvement because it reduces the pressure drop between the outlet of the diffuser and the upstream of the system. injection, by feeding these systems more directly. It also makes it possible to supply more symmetrically gas to the external and internal bypass zones of the chamber, and also to better supply gas to the internal bypass zone. In addition, the gas supply of the different intake channels of each injection system is also more homogeneous.

Les documents EP 1 788 310 et US 2007/0183890 décrivent d'autres exemples de diffuseurs.The documents EP 1 788 310 and US 2007/0183890 describe other examples of broadcasters.

L'invention vise à perfectionner ce dernier type de turbomachine, pour améliorer encore l'alimentation des systèmes d'injection.The invention aims to improve the latter type of turbomachine, to further improve the supply of injection systems.

Pour atteindre ce but, l'invention a pour objet une turbomachine, du dernier type précité (i.e. avec une partie aval du diffuseur inclinée par rapport à l'axe de la turbomachine, en direction du fond de chambre), selon l'objet de la revendication indépendante 1.To achieve this object, the subject of the invention is a turbomachine, of the last aforementioned type (ie with a downstream part of the diffuser inclined with respect to the axis of the turbomachine, in the direction of the chamber bottom), according to the object of the independent claim 1.

Lors des recherches ayant conduit à l'invention, les inventeurs ont trouvé que les aubes de redressement de la partie aval du diffuseur induisent des perturbations dans l'écoulement de gaz, en aval de leur bord de fuite (on parle de sillages), et que ces perturbations ont un effet néfaste sur l'alimentation en gaz des systèmes d'injection. Notamment, cette alimentation en gaz est moins symétrique autour de l'axe d'injection de ces systèmes. Or, plus on s'éloigne du bord de fuite des aubes de redressement, plus ces perturbations diminuent. L'invention propose donc d'éloigner suffisamment les orifices d'injection (et donc les systèmes d'injection) du bord de fuite des aubes de redressement, pour que ces orifices soient dans une zone de faible perturbation, voire de perturbation nulle, et ainsi améliorer l'alimentation des systèmes d'injection.During the research that led to the invention, the inventors have found that the straightening vanes of the downstream part of the diffuser induce disturbances in the gas flow downstream of their trailing edge (we speak of wakes), and these disturbances have a detrimental effect on the gas supply of the injection systems. In particular, this gas supply is less symmetrical around the injection axis of these systems. However, the further one gets away from the trailing edge of the straightening vanes, the more these disturbances diminish. The invention therefore proposes to distance sufficiently the injection orifices (and therefore the injection systems) from the trailing edge of the straightening vanes, so that these orifices are in an area of slight disturbance or even zero disturbance, and thus improve the supply of the injection systems.

Grâce à l'invention, on constate une meilleure alimentation du système d'injection (en particulier plus symétrique autour de l'axe d'injection), ce qui permet, notamment, d'améliorer la stabilité de la combustion à bas régime, de mieux maîtriser le profil des températures en sortie de chambre et de limiter le risque de couplage non-stationnaire entre l'écoulement en sortie de diffuseur et la combustion.Thanks to the invention, there is a better supply of the injection system (in particular more symmetrical about the injection axis), which allows, in particular, to improve the stability of the low-speed combustion, better control of the temperature profile at the outlet of the chamber and to limit the risk of non-stationary coupling between the outlet flow of the diffuser and the combustion.

D'un autre côté, les orifices d'injection ne doivent pas être trop éloignés des aubes de redressement pour limiter la perte de charge par éclatement entre la sortie du diffuseur et l'amont du système d'injection. Aussi, selon un mode de réalisation de l'invention, dans un plan de coupe contenant l'axe de la turbomachine et passant par le centre d'un desdits orifices d'injection, la distance, en abscisse curviligne, le long d'une ligne de courant entre le milieu de la veine de courant au bord de fuite des aubes de redressement et ledit centre, est inférieure ou égale à 9 fois la hauteur de ladite veine de courant au bord de fuite des aubes de redressement.On the other hand, the injection orifices must not be too far from the straightening vanes to limit the burst pressure drop between the outlet of the diffuser and the upstream of the injection system. Also, according to one embodiment of the invention, in a cutting plane containing the axis of the turbomachine and passing through the center of one of said injection orifices, the distance, in abscissa curvilinear, along a current line between the middle of the current stream at the trailing edge of the straightening vanes and said center, is less than or equal to 9 times the height of said stream of current at the trailing edge of the straightening vanes.

Pour améliorer le redressement de l'écoulement de gaz traversant la partie aval du diffuseur, il est préférable que le nombre d'aubes de redressement soit important sans toutefois générer un blocage aérodynamique préjudiciable à la marge au pompage de l'étage centrifuge. Aussi, selon un mode de réalisation de l'invention, le nombre d'aubes de redressement est supérieur au nombre de systèmes d'injection. Selon l'invention, le nombre d'aubes de redressement est au minimum 4 fois plus important que le nombre de systèmes d'injection. Plus le nombre d'aubes de redressement est important, plus les structures turbulentes induites par ces aubes sont nombreuses (mais de taille plus réduite), et plus l'invention présente d'intérêt, puisqu'elle permet de dissiper ces structure turbulentes sur la distance recommandée entre le bord de fuite des aubes de redressement et les orifices d'injection.To improve the recovery of the flow of gas passing through the downstream part of the diffuser, it is preferable for the number of straightening vanes to be large without, however, generating an aerodynamic blocking detrimental to the pumping margin of the centrifugal stage. Also, according to one embodiment of the invention, the number of straightening vanes is greater than the number of injection systems. According to the invention, the number of straightening vanes is at least 4 times greater than the number of injection systems. The greater the number of straightening vanes, the more turbulent structures induced by these vanes are numerous (but of smaller size), and the more the invention is of interest since it makes it possible to dissipate these turbulent structures on the recommended distance between the trailing edge of the straightening vanes and the injection ports.

L'invention et ses avantages seront mieux compris à la lecture de la description détaillée qui suit, d'exemples de réalisation de l'invention donnés à titre illustratif et non limitatif. Cette description fait référence aux figures annexées sur lesquelles :

  • les figures 1 à 3 représentent de manière schématique, en demi coupe axiale, le compresseur, le diffuseur et le module chambre de combustion de trois exemples de turboréacteur d'avion, conformes à l'invention;
  • la figure 4 représente le module chambre de combustion de la figure 1, et les lignes de courant de l'écoulement de gaz traversant ce module; et
  • la figure 5 représente partiellement la partie aval du diffuseur de la turbomachine de la figure 1, en section radiale, suivant le plan de coupe V-V.
The invention and its advantages will be better understood on reading the following detailed description of embodiments of the invention given for illustrative and non-limiting. This description refers to the appended figures in which:
  • the Figures 1 to 3 schematically represent, in half axial section, the compressor, the diffuser and the chamber module of combustion of three examples of aircraft turbojet, according to the invention;
  • the figure 4 represents the combustion chamber module of the figure 1 , and the current lines of the gas flow flowing through this module; and
  • the figure 5 partially represents the downstream part of the diffuser of the turbomachine of the figure 1 , in radial section, according to the VV cutting plane.

Les figures 1 à 3 représentent un exemple de turboréacteur, en demi-coupe axiale selon un plan de coupe contenant l'axe de rotation A du rotor du turboréacteur.The Figures 1 to 3 represent an example of a turbojet engine in axial half-section along a section plane containing the rotation axis A of the turbojet rotor.

Le gaz traversant le turboréacteur est de l'air.The gas passing through the turbojet is air.

Le turboréacteur comprend un compresseur 10 haute pression dont la partie avale (visible sur la figures 1 à 3) est composée d'un étage centrifuge, un diffuseur 20 annulaire raccordé en aval du compresseur 10, ce diffuseur débouchant dans un espace 30 entourant une chambre de combustion 40 annulaire. Cet espace 30 est délimité par un carter extérieur 32 et un carter intérieur 34 concentriques. La chambre de combustion 40 est soutenue par des brides de fixation reliées aux carters 32 et 34.The turbojet engine comprises a high pressure compressor 10 whose part swallows (visible on the Figures 1 to 3 ) is composed of a centrifugal stage, an annular diffuser connected downstream of the compressor 10, this diffuser opening into a space 30 surrounding an annular combustion chamber 40. This space 30 is delimited by an outer casing 32 and a concentric inner casing 34. The combustion chamber 40 is supported by fixing clamps connected to the housings 32 and 34.

Le compresseur 10 centrifuge comprend un impulseur rotatif 12. Cet impulseur 12 comprend une série d'aubes 14 mobiles, entraînées en rotation. L'impulseur 12 est réalisé de façon à accélérer l'air qui le traverse et, de ce fait, à accroître l'énergie cinétique de cet air.The centrifugal compressor comprises a rotary impeller 12. This impeller 12 comprises a series of moving blades 14, driven in rotation. The impeller 12 is designed to accelerate the air passing through it and thereby increase the kinetic energy of this air.

Le diffuseur 20 présente un espace annulaire entourant l'impulseur. Le diffuseur 20 sert à réduire la vitesse de l'air quittant l'impulseur et, de ce fait, à accroître sa pression statique. Le diffuseur 20 des figures est du type à aubes.The diffuser 20 has an annular space surrounding the impeller. The diffuser 20 serves to reduce the speed of the air leaving the impeller and thereby increase its static pressure. The diffuser 20 of the figures is of the paddle type.

Ce diffuseur 20 a une partie amont 21 annulaire orientée radialement qui présente une série de passages de diffusion 22 raccordés à la sortie du compresseur 10 pour récupérer l'air accéléré sortant de l'impulseur 12. Ces passages de diffusion 22 ont une section qui s'accroît progressivement de l'amont vers l'aval afin de diffuser le courant d'air sortant de l'impulseur. Ces passages de diffusion 22 sont formés par une série d'aubes 23 espacées circulairement. Au niveau de l'entrée de la partie amont 21, ces aubes 23 sont proches les unes des autres. Ces aubes 23 s'éloignent circonférentiellement les unes des autres à mesure qu'elles se rapprochent de la sortie de la partie amont 21.This diffuser 20 has a radially oriented annular upstream portion 21 which has a series of diffusion passages 22 connected to the output of the compressor 10 to recover the accelerated air coming out of the impeller 12. These diffusion passages 22 have a section which progressively increases from upstream to downstream in order to diffuse the air flow leaving the impeller. These diffusion passages 22 are formed by a series of blades 23 spaced apart in a circular manner. At the inlet of the upstream portion 21, these blades 23 are close to each other. These vanes 23 circumferentially move away from each other as they approach the outlet of the upstream portion 21.

En aval de la partie amont 21, le diffuseur 20 comprend une partie intermédiaire 24 annulaire coudée, pour courber le trajet d'écoulement du diffuseur et ramener l'écoulement d'air en direction de la chambre de combustion 40.Downstream of the upstream portion 21, the diffuser 20 comprises an annular intermediate portion 24 bent, to curve the flow path of the diffuser and reduce the flow of air towards the combustion chamber 40.

En aval de cette partie intermédiaire 24, le diffuseur 20 comprend une partie aval 25 annulaire comprenant une série d'aubes de redressement 26 circulairement espacées, pour supprimer ou limiter le tourbillonnement circonférentiel de l'écoulement d'air sortant des passages de diffusion 22, avant que cet écoulement d'air n'entre dans l'espace 30. La figure 5 représente les aubes de redressement 26 en section, dans un plan de coupe perpendiculaire à l'axe moyen X de la veine de courant en sortie de la partie aval 25 du diffuseur. Comme on le voit sur cette figure, les aubes de redressement 26 s'étendent radialement entre les parois intérieure et extérieure de la partie aval 25 du diffuseur.Downstream of this intermediate portion 24, the diffuser 20 comprises an annular downstream portion 25 comprising a series of circumferentially spaced rectifying vanes 26 for eliminating or limiting the circumferential swirling of the outgoing air flow from the diffusion passages 22, before this airflow enters the space 30. The figure 5 represents the straightening vanes 26 in section, in a sectional plane perpendicular to the mean axis X of the stream of current at the outlet of the downstream part 25 of the diffuser. As seen in this figure, the straightening vanes 26 extend radially between the inner and outer walls of the downstream part 25 of the diffuser.

La chambre de combustion 40 comprend une paroi intérieure 42 annulaire, une paroi extérieure 43 annulaire et un fond de chambre 41 annulaire disposé entre lesdites parois 42, 43, dans la région amont de ladite chambre. Ce fond de chambre 41 présente des orifices d'injection 44 répartis circulairement autour de l'axe A. Des systèmes d'injection 45 sont montés sur le fond de chambre, au travers desdits orifices d'injection 44 (à raison d'un système d'injection 45 par orifice d'injection 44). Ces systèmes d'injection 45 permettent d'injecter le mélange air/carburant qui est brûlé dans la chambre de combustion 40. Le carburant de ce mélange est amené jusqu'aux systèmes d'injection 45 par une conduite 46 d'alimentation en carburant traversant l'espace 30.The combustion chamber 40 comprises an annular inner wall 42, an annular outer wall 43 and an annular chamber bottom 41 disposed between said walls 42, 43, in the upstream region of said chamber. This chamber bottom 41 has injection orifices 44 distributed circumferentially about the axis A. Injection systems 45 are mounted on the chamber bottom, through said injection ports 44 (due to a system injection 45 by injection port 44). These injection systems 45 make it possible to inject the air / fuel mixture which is burned into the combustion chamber 40. The fuel of this mixture is fed to the injection systems 45 via a fuel feed pipe 46 running through the space 30.

On appelle communément module de chambre de combustion l'ensemble constitué par la chambre de combustion 40, ses carters et son environnement proche.The combustion chamber 40, its housings and its close environment are commonly referred to as the combustion chamber module.

La chambre de combustion 40 est inclinée par rapport à l'axe A du turboréacteur d'un angle aigu B (non nul). Plus cet angle aigu B est grand et plus l'encombrement axial du module chambre de combustion est réduit.The combustion chamber 40 is inclined relative to the axis A of the turbojet engine of an acute angle B (non-zero). The higher this acute angle B is, the smaller the axial size of the combustion chamber module.

Afin que l'écoulement d'air sortant du diffuseur 20 soit orienté en direction des systèmes d'injection 45, la partie aval 25 du diffuseur 20 est inclinée par rapport à l'axe A du turboréacteur, en direction de la chambre de combustion, de sorte que, dans un plan de coupe contenant l'axe du turboréacteur, l'axe moyen X de la veine de courant en sortie de la partie aval 25 du diffuseur, passe par le fond de chambre 41, entre le rayon maximum R et le rayon minimum r du fond de chambre 41. Le fond de chambre 41 étant annulaire et centré sur l'axe A, les rayons r et R s'étendent depuis l'axe A suivant une direction radiale. Pour illustrer ceci, les figures 1 à 3 représentent trois exemples de chambre de combustion 40 pour lesquelles le critère d'inclinaison précité est respecté. Sur la figure 2, l'axe moyen X de la veine de courant en sortie de la partie aval 25 du diffuseur, passe par le fond de chambre 41 à proximité du rayon maximum R. Sur la figure 3, l'axe moyen X de la veine de courant en sortie de la partie aval 25 du diffuseur, passe par le fond de chambre 41 à proximité du rayon minimum r. Sur la figure 1, l'axe moyen X de la veine de courant en sortie de la partie aval 25 du diffuseur, passe par le fond de chambre 41 au niveau du centre C d'un orifice d'injection 44.So that the flow of air leaving the diffuser 20 is oriented towards the injection systems 45, the downstream portion 25 of the diffuser 20 is inclined relative to the axis A of the turbojet engine, in the direction of the combustion chamber, so that, in a section plane containing the axis of the turbojet, the average axis X of the stream of current at the outlet of the downstream part 25 of the diffuser, passes through the chamber bottom 41, between the maximum radius R and the minimum radius r of the chamber bottom 41. The chamber bottom 41 being annular and centered on the axis A, the r and r rays extend from the axis A in a radial direction. To illustrate this, Figures 1 to 3 represent three examples of combustion chamber 40 for which the aforementioned inclination criterion is respected. On the figure 2 the mean axis X of the stream of current at the outlet of the downstream part 25 of the diffuser passes through the chamber bottom 41 close to the maximum radius R. On the figure 3 the mean axis X of the current stream at the outlet of the downstream part 25 of the diffuser passes through the chamber bottom 41 near the minimum radius r. On the figure 1 the mean axis X of the stream of current at the outlet of the downstream part 25 of the diffuser passes through the chamber bottom 41 at the center C of an injection orifice 44.

Conformément à l'invention et en référence à la figure 4 qui représente le module chambre de combustion de la figure 1 et les lignes de courant de l'écoulement d'air traversant ce module, si on se place dans le plan de coupe des figures 1 et 4 qui contient l'axe A du turboréacteur et qui passe par le centre C d'un orifice d'injection 44, la distance, en abscisse curviligne, le long d'une ligne de courant L entre le milieu O de la veine de courant au bord de fuite des aubes de redressement 26 et ledit centre C (c'est-à-dire la longueur de la portion de courbe appartenant à la ligne de courant L et allant de O à C), est supérieure ou égale à 3 fois la hauteur h de ladite veine de courant au bord de fuite des aubes de redressement (cette hauteur h correspond à la hauteur des aubes de redressement 26). D'autre part, cette distance, en abscisse curviligne, est inférieure ou égale à 9 fois la hauteur de ladite veine de courant au bord de fuite des aubes de redressement.In accordance with the invention and with reference to figure 4 which represents the combustion chamber module of the figure 1 and the current lines of the air flow passing through this module, if one places oneself in the plane of section of the figures 1 and 4 which contains the axis A of the turbojet and which passes through the center C of an injection orifice 44, the distance, in curvilinear abscissa, along a stream line L between the middle O of the stream of current at trailing edge of the straightening vanes 26 and said center C (that is to say the length of the portion of the curve belonging to the current line L and ranging from 0 to C), is greater than or equal to 3 times the height h of said stream vein at the trailing edge of the straightening vanes (this height h corresponds to the height of the straightening vanes 26). On the other hand, this distance, curvilinear abscissa, is less than or equal to 9 times the height of said stream of current at the trailing edge of the straightening vanes.

Claims (2)

  1. A turbomachine comprising
    • an annular combustion chamber (40) with a chamber end wall (41) presenting injection orifices (44) having injection systems (45) mounted therein;
    • a compressor (10) having a centrifugal downstream stage; and
    • an annular diffuser (20) enabling the gas flow leaving the compressor (10) to be diffused and said flow of gas to be directed towards said injection systems, the diffuser comprising: a radially-oriented upstream portion (21) that presents diffusion passages (22) connected to the outlet of the compressor; an elbow-shaped intermediate portion (24); and a downstream portion (25) comprising a series of circularly spaced apart flow-straightening vanes (26), the downstream portion (25) being inclined relative to the axis (A) of the turbomachine towards the combustion chamber, in such a manner that in the section plane containing the axis (A) of the turbomachine, the mean axis (X) of the flow path at the outlet from the downstream portion (25) of the diffuser (20) passes through the chamber end wall (41) between the maximum radius (R) and the minimum radius (r) of the chamber end wall (41);
    the turbomachine being characterized in that, in the section plane containing the axis (A) of the turbomachine and passing via the center (C) of one of said injection orifices (44), the curvilinear abscissa distance along a flow line (12) between the middle (0) of the flow path at the trailing edges of the flow-straightening vanes (26) and said center (C), is greater than or equal to three times the height (h) of said flow path at the trailing edges of the flow-straightening vanes (26), and in that the number of flow-straightening vanes (26) is at least four times greater than the number of injection systems (45).
  2. The turbomachine according to claim 1, in which, in a section plane containing the axis (A) of the turbomachine and passing via the center (C) of one of said injection orifices (44), the curvilinear abscissa distance along a flow line (L) between the middle (O) of the flow path at the trailing edges of the flow-straightening vanes (26) and said center (C) is less than or equal to nine times the height (h) of said flow path at the trailing edges of the flow-straightening vanes (26).
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FR2904048B1 (en) 2006-07-19 2012-12-14 Snecma COMBUSTION CHAMBER WALL VENTILATION SYSTEM IN TURBOMACHINE
FR2905166B1 (en) 2006-08-28 2008-11-14 Snecma Sa ANNULAR COMBUSTION CHAMBER OF A TURBOMACHINE.
FR2906350B1 (en) 2006-09-22 2009-03-20 Snecma Sa ANNULAR COMBUSTION CHAMBER OF A TURBOMACHINE
FR2906296A1 (en) * 2006-09-26 2008-03-28 Snecma Sa DEVICE FOR FASTENING A FIXED BLADE IN AN ANNULAR CASE FOR TURBOMACHINE, TURBOREACTOR INCORPORATING THE DEVICE AND METHOD FOR MOUNTING THE BLADE.
FR2910115B1 (en) 2006-12-19 2012-11-16 Snecma DEFLECTOR FOR BOTTOM OF COMBUSTION CHAMBER, COMBUSTION CHAMBER WHERE IT IS EQUIPPED AND TURBOREACTOR COMPRISING THEM
FR2911669B1 (en) 2007-01-23 2011-09-16 Snecma FURNITURE FOR COMBUSTION CHAMBER, COMBUSTION CHAMBER WHEN EQUIPPED AND TURBOREACTOR COMPRISING THEM.
FR2920033B1 (en) * 2007-08-13 2014-08-22 Snecma TURBOMACHINE WITH DIFFUSER
FR2922630B1 (en) 2007-10-22 2015-11-13 Snecma COMBUSTION CHAMBER WALL WITH OPTIMIZED DILUTION AND COOLING, COMBUSTION CHAMBER AND TURBOMACHINE WHILE ENHANCED
FR2922629B1 (en) 2007-10-22 2009-12-25 Snecma COMBUSTION CHAMBER WITH OPTIMIZED DILUTION AND TURBOMACHINE WHILE MUNIED

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US20090047127A1 (en) 2009-02-19
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CN101368512B (en) 2013-06-12
EP2071152A2 (en) 2009-06-17
CA2638817A1 (en) 2009-02-13
FR2920032A1 (en) 2009-02-20
EP2071152A3 (en) 2017-08-02
FR2920032B1 (en) 2014-08-22
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CN101368512A (en) 2009-02-18
US8047777B2 (en) 2011-11-01
RU2008133234A (en) 2010-02-20
CA2638817C (en) 2016-04-12

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